U.S. patent number 10,589,635 [Application Number 16/289,900] was granted by the patent office on 2020-03-17 for active voltage control for hybrid electric aircraft.
This patent grant is currently assigned to The Boeing Company. The grantee listed for this patent is The Boeing Company. Invention is credited to Kamiar J. Karimi, Mark E. Liffring, Eugene V. Solodovnik.
United States Patent |
10,589,635 |
Solodovnik , et al. |
March 17, 2020 |
Active voltage control for hybrid electric aircraft
Abstract
A solid-state high-voltage direct-current (HVDC) bus voltage
controller to provide active power flow control in a hybrid
electric aircraft power supply system. The HVDC bus voltage
controller includes an active voltage controller and an active
rectifier unit configured to control the HVDC bus voltage using the
PWM control technique. In one implementation, the active rectifier
unit includes high-power and high-frequency semiconductor switches
with fast turn-off capabilities. The active voltage controller
sends an HVDC bus reference voltage to the active rectifier unit.
The low-level controller inside the active rectifier unit is
configured to control the HVDC bus voltage to match the HVDC bus
reference voltage.
Inventors: |
Solodovnik; Eugene V. (Lake
Stevens, WA), Liffring; Mark E. (Seattle, WA), Karimi;
Kamiar J. (Kirkland, WA) |
Applicant: |
Name |
City |
State |
Country |
Type |
The Boeing Company |
Chicago |
IL |
US |
|
|
Assignee: |
The Boeing Company (Chicago,
IL)
|
Family
ID: |
69630166 |
Appl.
No.: |
16/289,900 |
Filed: |
March 1, 2019 |
Current U.S.
Class: |
1/1 |
Current CPC
Class: |
B60L
58/20 (20190201); B60L 50/13 (20190201); H02J
7/1415 (20130101); B60R 16/033 (20130101); H02J
7/1492 (20130101); B60L 15/2045 (20130101); H02J
2310/44 (20200101) |
Current International
Class: |
B60L
50/13 (20190101); B60R 16/033 (20060101); B60L
58/20 (20190101); B60L 15/20 (20060101) |
References Cited
[Referenced By]
U.S. Patent Documents
Primary Examiner: Berhane; Adolf D
Attorney, Agent or Firm: Ostrager Chong Flaherty &
Broitman P.C.
Claims
The invention claimed is:
1. A hybrid electrical power supply system for supplying current to
a load, the hybrid electrical power supply system comprising: an AC
power source; an AC bus; first switching means arranged to
switchably couple AC power from the AC power source to the AC bus;
an HVDC power source; an active voltage controller configured to
generate a control signal representing a magnitude of a HVDC bus
reference voltage; an active rectifier unit comprising an active
rectifier configured to convert AC power to HVDC power and a
controller configured to control the active rectifier to adjust an
HVDC bus voltage in dependence on the control signal received from
the active voltage controller; second switching means arranged to
switchably couple AC power from the AC bus to the active rectifier
unit; an HVDC bus coupled to receive DC power from one or both of
the HVDC power source and active rectifier unit in dependence on
the HVDC bus voltage from the active rectifier unit; and an HVDC
load coupled to receive HVDC power from the HVDC bus.
2. The hybrid electrical power supply system as recited in claim 1,
wherein the HVDC load comprises a motor controller coupled to
receive HVDC power from the HVDC bus.
3. The hybrid electrical power supply system as recited in claim 1,
further comprising an AC load and third switching means arranged to
switchably couple AC power from the AC bus to the AC load.
4. The hybrid electrical power supply system as recited in claim 1,
wherein: the active voltage controller comprises a computer or
processor programmed to generate the control signal in different
forms corresponding to different magnitudes of the HVDC bus
reference voltage; and the controller of the active rectifier unit
is configured to control the active rectifier to adjust the HVDC
bus voltage of the HVDC bus in dependence on the magnitude of the
HVDC bus reference voltage.
5. The hybrid electrical power supply system as recited in claim 1,
wherein the AC power source comprises an electrical generator and
the HVDC power source comprises a battery.
6. The hybrid electrical power supply system as recited in claim 5,
wherein the control signal is a command to increase the HVDC bus
voltage when current out of the battery is greater than a maximum
discharge rate.
7. The hybrid electrical power supply system as recited in claim 5,
wherein the control signal is a command to decrease the HVDC bus
voltage when current into the battery is greater than a maximum
charge rate.
8. The hybrid electrical power supply system as recited in claim 5,
further comprising a flight power management controller configured
to send power flow management commands to the active voltage
controller to optimize overall energy usage and reduce fuel burn
during a particular mission of an aircraft.
9. The hybrid electrical power supply system as recited in claim 8,
wherein the active voltage controller generates an initiate descent
message when both of first and second conditions are met, the first
condition being that a state of charge of the battery is less than
a minimum state of charge reserve, and the second condition being
that a fuel reserve is less than a minimum fuel reserve.
10. The hybrid electrical power supply system as recited in claim
8, wherein the flight power management controller is further
configured to send a command to the active voltage controller to
charge the battery if the aircraft is on ground, the battery is
discharged and a power source is available.
11. The hybrid electrical power supply system as recited in claim
8, wherein the flight power management controller is further
configured to perform operations comprising: determining what is a
current phase of the flight of the aircraft; determining a battery
state of charge; determining a fuel level; adjusting a battery
usage profile in dependence on the flight phase, battery state of
charge and fuel level; setting a battery discharging or charging
current in accordance with the adjusted battery usage profile; and
commanding the active voltage controller to cause the battery to be
discharged or charged in accordance with the set battery
discharging or charging current.
12. An aircraft comprising: an AC power source; an AC bus; first
switching means arranged to switchably couple AC power from the AC
power source to the AC bus; an HVDC power source; an active voltage
controller configured to generate a control signal representing a
magnitude of a HVDC bus reference voltage based on a current of the
HVDC power source; an active rectifier unit comprising an active
rectifier configured to convert AC power to HVDC power and a
controller configured to control the active rectifier to adjust an
HVDC bus voltage in dependence on the control signal received from
the active voltage controller; second switching means arranged to
switchably couple AC power from the AC bus to the active rectifier
unit; an HVDC bus coupled to receive DC power from one or both of
the HVDC power source and active rectifier unit in dependence on
the HVDC bus voltage from the active rectifier unit; an electric
propulsion motor coupled to receive HVDC power from the HVDC bus;
and a propeller operatively coupled to the electric propulsion
motor.
13. The aircraft as recited in claim 12, wherein: the active
voltage controller comprises a computer or processor programmed to
generate the control signal in different forms corresponding to
different magnitudes of the HVDC bus reference voltage; and the
controller of the active rectifier unit is configured to control
the active rectifier to adjust the HVDC bus voltage of the HVDC bus
in dependence on the magnitude of the HVDC bus reference
voltage.
14. The aircraft as recited in claim 12, further comprising a
flight power management controller configured to send power flow
management commands to the active voltage controller to optimize
overall energy usage and reduce fuel burn during a particular
mission.
15. The aircraft as recited in claim 12, wherein the AC power
source comprises an electrical generator and the HVDC power source
comprises a battery.
16. The aircraft as recited in claim 15, wherein the active voltage
controller generates an initiate descent message when both of first
and second conditions are met, the first condition being that a
state of charge of the battery is less than a minimum state of
charge reserve, and the second condition being that a fuel reserve
is less than a minimum fuel reserve.
17. The aircraft as recited in claim 15, further comprising a
flight power management controller configured to send a command to
the active voltage controller to charge the battery if the aircraft
is on ground, the battery is discharged and a power source is
available.
18. The aircraft as recited in claim 15, further comprising a
flight power management controller configured to perform
operations, if the aircraft is in flight, comprising: determining
what is a current phase of the flight of the aircraft; determining
a battery state of charge; determining a fuel level; adjusting a
battery usage profile in dependence on the flight phase, battery
state of charge and fuel level; setting a battery discharging or
charging current in accordance with the adjusted battery usage
profile; and commanding the active voltage controller to cause the
battery to be discharged or charged in accordance with the set
battery discharging or charging current.
19. A method for supplying HVDC power to an HVDC load via an HVDC
bus that is coupled to receive DC power from one or both of an HVDC
power source and an active rectifier unit that is coupled to an AC
power source, the method comprising: generating a control signal in
an active voltage controller in a form representing a magnitude of
a HVDC bus reference voltage; and controlling the active rectifier
unit to adjust an HVDC bus voltage of the HVDC bus in dependence on
the magnitude of the HVDC bus reference voltage.
20. The method as recited in claim 19, wherein the AC and HVDC
power sources are respectively an electric generator and a battery
onboard an aircraft, further comprising: sending power flow
management commands from a flight power management controller
onboard the aircraft to the active voltage controller to optimize
overall energy usage and reduce fuel burn during a particular
mission of an aircraft by adjusting the magnitude of the HVDC bus
reference voltage relative to a voltage of the battery.
Description
BACKGROUND
This disclosure generally relates to hybrid electrical power
sources having two or more electrical energy sources that supply
energy to a connected load. In particular, the technology disclosed
herein relates to hybrid electrical power sources comprising one or
more batteries and one or more electric generators driven by
internal combustion engines or gas turbine engines.
Some aircraft have electrically powered propulsion systems
(hereinafter "electric aircraft"). In such aircraft, electric
motors convert electrical power into mechanical power for use by
the propulsion system. For example, an electric motor may turn one
or more propellers on the aircraft to provide thrust. An electric
aircraft may take various forms. For example, the electric aircraft
may be an aircraft, a rotorcraft, a helicopter, a quadcopter, an
unmanned aerial vehicle, or some other suitable type of
aircraft.
When electric motors are used for propulsion of the aircraft,
electrical energy may be supplied by a power source. For instance,
electrical energy may be supplied using a battery system. The load
on the battery system or other power source is an important
consideration for the design and manufacturing of the aircraft. For
example, the amount of electrical energy used by the electric motor
to move the aircraft during various stages of flight may be
important. Electric motors that use battery systems may require the
battery to be recharged after a specified amount of time, distance,
electrical energy use, or a combination thereof.
Some electric aircraft have a hybrid electric power architecture
(hereinafter "hybrid electric aircraft") in which at least two
different types of power sources are connected in parallel to a
load. The electrical energy sources will often have different
electrical characteristics. For example, the electrical energy
sources may be a battery and an electric generator driven by an
internal combustion engine or a gas turbine engine. The battery
supplies electrical power to an electric motor that is arranged to
convert electrical power into mechanical power for use by the
propulsion system of the aircraft.
In the case of a battery-equipped hybrid electric aircraft, the
battery voltage cannot be actively controlled. Battery voltage is
determined by the number of cells, type of cells, battery state of
charge (SOC), loading and other factors. It is necessary to control
power flow to and from the battery. Rate of charge or discharge of
the battery is important and should be controlled to avoid thermal
runaway.
For hybrid electric aircraft, the batteries are large and designed
to provide a large amount of power for the purpose of propulsion.
The batteries are paralleled with other power sources, such as
electric generators. In one implementation, the battery is
connected to a high-voltage direct-current (HVDC) bus, which is
also supplied by the generator source(s). As used in the aerospace
industry and herein, the term "high voltage" in the context of
direct current means any DC voltage higher than 500 V.sub.DC. Such
DC high voltage is typically derived from rectification of
three-phase 230 V.sub.AC power.
There are no existing solutions for active power flow control and
battery power management control for hybrid electric aircraft that
employ batteries with hundreds of kilowatt-hours of energy. A
system and method for tightly controlling the power flow to and
from the battery at the HVDC connection is wanted.
SUMMARY
The subject matter disclosed in some detail below is directed to a
power supply system architecture for a hybrid electric aircraft.
The power supply system includes a solid-state HVDC bus voltage
controller that provides active control of the flow of DC power to
the propulsion motors of the hybrid electric aircraft power supply
system. The HVDC bus voltage controller includes an active voltage
controller and an active rectifier unit configured to control the
HVDC bus voltage using a pulse-width modulated (PWM) control
technique. In one implementation, the active rectifier unit
includes high-power and high-frequency semiconductor switches with
fast turn-off capabilities. The PWM control technique is used to
effect a desired power transmission. Control of the voltage of the
HVDC bus enables controlled charge or discharge of the battery and
desirable power flow in the system.
More specifically, the power supply system includes an
alternating-current (AC) generator connected to an AC bus, a
high-voltage direct-current (HVDC) battery connected to an HVDC
bus, an AC-to-DC power converter connecting the AC bus to the HVDC
bus, and an active voltage controller configured to control the
power flow to and from the battery at an HVDC connection. The
active voltage controller is situated at an output of the battery.
The active voltage controller sends a control signal representing
the magnitude of an HVDC bus reference voltage to the active
rectifier unit. The low-level controller inside the active
rectifier unit is configured to control the HVDC bus voltage to
match the HVDC bus reference voltage.
The active voltage control method disclosed herein allows for
accurate control of the battery charge or discharge rate. In
addition, the active voltage controller allows for precise power
flow management in the hybrid electric aircraft. Depending on
flight phase, aircraft mission, battery SOC, remaining hydrocarbon
fuel, flight conditions, etc., the active voltage controller is
configured to accurately manage the power flow to the propulsion
system, in some instances taking more power from the battery source
and less power from the conventional engine source or in other
instances preserving battery power, while using more power
generated from an engine source. The architecture disclosed herein
is reliable and low weight. The use of an active voltage controller
could also extend the life of the batteries, resulting in cost
savings.
Although various embodiments of systems and methods for actively
controlling the voltage of a bus that supplies electric power from
a battery and/or generator to propulsion motors of a hybrid
electric aircraft will be described in some detail below, one or
more of those embodiments may be characterized by one or more of
the following aspects.
One aspect of the subject matter disclosed in detail below is a
hybrid electrical power supply system for supplying current to a
load, the hybrid electrical power supply system comprising: an AC
power source; an AC bus; first switching means arranged to
switchably couple AC power from the AC power source to the AC bus;
an HVDC power source; an active voltage controller configured to
generate a control signal representing a magnitude of a HVDC bus
reference voltage; an active rectifier unit comprising an active
rectifier configured to convert AC power to HVDC power and a
controller configured to control the active rectifier to adjust an
HVDC bus voltage in dependence on the control signal received from
the active voltage controller; second switching means arranged to
switchably couple AC power from the AC bus to the active rectifier
unit; an HVDC bus coupled to receive DC power from one or both of
the HVDC power source and active rectifier unit in dependence on
the HVDC bus voltage from the active rectifier unit; and an HVDC
load coupled to receive HVDC power from the HVDC bus. The active
voltage controller comprises a computer or processor programmed to
generate the control signal in different forms corresponding to
different magnitudes of the HVDC bus reference voltage; and the
controller of the active rectifier unit is configured to control
the active rectifier to adjust the HVDC bus voltage of the HVDC bus
in dependence on the magnitude of the HVDC bus reference
voltage.
In accordance with some embodiments of the hybrid electrical power
supply system described in the immediately preceding paragraph, the
HVDC load comprises a motor controller coupled to receive HVDC
power from the HVDC bus, the AC power source comprises an
electrical generator and the HVDC power source comprises a
battery.
In accordance with some embodiments, the hybrid electrical power
supply system further comprises a flight power management
controller configured to send power flow management commands to the
active voltage controller to optimize overall energy usage and
reduce fuel burn during a particular mission of an aircraft. The
active voltage controller generates an initiate descent message
when both of first and second conditions are met, the first
condition being that a state of charge of the battery is less than
a minimum state of charge reserve, and the second condition being
that a fuel reserve is less than a minimum fuel reserve. The flight
power management controller is further configured to send a command
to the active voltage controller to charge the battery if the
aircraft is on ground, the battery is discharged and a power source
is available. In addition, the flight power management controller
is further configured to perform operations comprising: determining
what is the current phase of the flight of the aircraft;
determining the battery state of charge; determining the fuel
level; adjusting a battery usage profile in dependence on the
flight phase, battery state of charge and fuel level; setting a
battery discharging or charging current in accordance with the
adjusted battery usage profile; and commanding the active voltage
controller to cause the battery to be discharged or charged in
accordance with the set battery discharging or charging
current.
Another aspect of the subject matter disclosed in detail below is
an aircraft comprising: an AC power source; an AC bus; first
switching means arranged to switchably couple AC power from the AC
power source to the AC bus; an HVDC power source; an active voltage
controller configured to generate a control signal representing a
magnitude of a HVDC bus reference voltage based on a current of the
HVDC power source; an active rectifier unit comprising an active
rectifier configured to convert AC power to HVDC power and a
controller configured to control the active rectifier to adjust an
HVDC bus voltage of the HVDC bus in dependence on the control
signal received from the active voltage controller; second
switching means arranged to switchably couple AC power from the AC
bus to the active rectifier unit; an HVDC bus coupled to receive DC
power from one or both of the HVDC power source and active
rectifier unit in dependence on the HVDC bus voltage from the
active rectifier unit; an electric propulsion motor coupled to
receive HVDC power from the HVDC bus; and a propeller operatively
coupled to the electric propulsion motor. The active voltage
controller comprises a computer or processor programmed to generate
the control signal in different forms corresponding to different
magnitudes of the HVDC bus reference voltage. The controller of the
active rectifier unit is configured to control the active rectifier
to adjust the HVDC bus voltage of the HVDC bus in dependence on the
magnitude of the HVDC bus reference voltage. The aircraft further
comprises a flight power management controller configured to send
power flow management commands to the active voltage controller to
optimize overall energy usage and reduce fuel burn during a
particular mission. In accordance with one embodiment, the AC power
source comprises an electrical generator and the HVDC power source
comprises a battery.
In accordance with one embodiment of the aircraft described in the
immediately preceding paragraph, the flight power management
controller is configured to send a command to the active voltage
controller to charge the battery if the aircraft is on ground, the
battery is discharged and a power source is available. In
accordance with the same or an alternative embodiment, the flight
power management controller is configured to perform operations, if
the aircraft is in flight, comprising: determining what is the
current phase of the flight of the aircraft; determining the
battery state of charge; determining the fuel level; adjusting a
battery usage profile in dependence on the flight phase, battery
state of charge and fuel level; setting a battery discharging or
charging current in accordance with the adjusted battery usage
profile; and commanding the active voltage controller to cause the
battery to be discharged or charged in accordance with the set
battery discharging or charging current.
A further aspect of the subject matter disclosed in detail below is
a method for supplying HVDC power to an HVDC load via an HVDC bus
that is coupled to receive DC power from one or both of an HVDC
power source and an active rectifier unit that is coupled to an AC
power source, the method comprising: generating a control signal in
an active voltage controller in a form representing a magnitude of
a HVDC bus reference voltage; and controlling the active rectifier
unit to adjust an HVDC bus voltage of the HVDC bus in dependence on
the magnitude of the HVDC bus reference voltage.
In accordance with one embodiment of the method described in the
immediately preceding paragraph, the AC and HVDC power sources are
respectively an electric generator and a battery onboard an
aircraft, and the method further comprises: sending power flow
management commands from a flight power management controller
onboard the aircraft to the active voltage controller to optimize
overall energy usage and reduce fuel burn during a particular
mission of an aircraft by adjusting the magnitude of the HVDC bus
reference voltage relative to a voltage of the battery.
Other aspects of systems and methods for actively controlling the
voltage of a bus that supplies electric power from a battery and/or
generator to propulsion motors of a hybrid electric aircraft are
disclosed below.
BRIEF DESCRIPTION OF THE DRAWINGS
The features, functions and advantages discussed in the preceding
section may be achieved independently in various embodiments or may
be combined in yet other embodiments. Various embodiments will be
hereinafter described with reference to drawings for the purpose of
illustrating the above-described and other aspects. None of the
diagrams briefly described in this section are drawn to scale.
FIG. 1 is a block diagram depicting a hybrid electric aircraft
power supply system architecture with a flight power management
controller (FPMC), at least one active voltage controller (AVC) and
at least one active rectifier unit (ARU) in accordance with one
embodiment.
FIG. 2 is a block diagram identifying some components of the
solid-state active rectifier unit included in the architecture
depicted in FIG. 1.
FIG. 3 is a flowchart identifying steps of a control algorithm
performed by the active voltage controller in accordance with one
proposed implementation.
FIG. 4 is a flowchart identifying steps of a control algorithm
performed by the flight power management controller in accordance
with one proposed implementation.
Reference will hereinafter be made to the drawings in which similar
elements in different drawings bear the same reference
numerals.
DETAILED DESCRIPTION
Illustrative embodiments of systems and methods for actively
controlling the voltage of a bus that supplies electric power from
a battery and/or generator to propulsion motors of a hybrid
electric aircraft are described in some detail below. However, not
all features of an actual implementation are described in this
specification. A person skilled in the art will appreciate that in
the development of any such embodiment, numerous
implementation-specific decisions must be made to achieve the
developer's specific goals, such as compliance with system-related
and business-related constraints, which will vary from one
implementation to another. Moreover, it will be appreciated that
such a development effort might be complex and time-consuming, but
would nevertheless be a routine undertaking for those of ordinary
skill in the art having the benefit of this disclosure.
A single hybrid electric aircraft may have one or more HVDC power
sources (e.g., batteries) and one or more generator sources driven
by a prime mover (e.g., a gas turbine engine or an internal
combustion engine). The multiple sources need to be controlled and
managed to achieve the optimal power extraction for the specific
mission given changing operating conditions of the batteries,
aircraft and environment.
More generally, the hybrid electrical power supply system proposed
herein is not limited in its application to hybrid electric
aircraft. In accordance with one architecture, the hybrid
electrical power supply system proposed herein includes the
following components: an AC power source; an AC bus; first
switching means arranged to switchably couple AC power from the AC
power source to the AC bus; an HVDC power source; an active voltage
controller configured to generate a control signal representing a
magnitude of a HVDC bus reference voltage; an active rectifier unit
comprising an active rectifier configured to convert AC power to
HVDC power and a controller configured to control the active
rectifier to adjust an HVDC bus voltage in dependence on the
control signal received from the active voltage controller; second
switching means arranged to switchably couple AC power from the AC
bus to the active rectifier unit; an HVDC bus coupled to receive DC
power from one or both of the HVDC power source and active
rectifier unit in dependence on the HVDC bus voltage from the
active rectifier unit; and an HVDC load coupled to receive HVDC
power from the HVDC bus. One example of an embodiment of the
foregoing architecture, designed for powering an electric aircraft,
will be described in some detail below.
FIG. 1 is a block diagram depicting in some detail a proposed
architecture for a power supply system 10 of a hybrid electric
aircraft in accordance with one embodiment. In this example, the
hybrid electric aircraft has four propellers 24 (two on each wing).
Only two propellers 24 rotatably mounted to one wing and associated
power supply system components are depicted in FIG. 1. The power
supply system 10 depicted in FIG. 1 further includes a flight power
management controller 12 (hereinafter "FPMC 12"). All of the other
components identified in FIG. 1 are associated with one wing of the
hybrid electric aircraft and form one half of the power supply
system 10. The other half (not shown in FIG. 1) of the power supply
system 10 includes duplicate components associated with the other
wing. The FPMC 12 manages the electrical propulsion power supplied
to both wings.
In accordance with one embodiment, each propeller 24 is driven to
rotate by a respective electric propulsion motor 26. In an
alternative embodiment, each electric propulsion motor 26 drives
multiple propellers. Each electric propulsion motor 26 operates
under the control of a respective propulsion motor controller 28,
which receives aircraft propulsion power from an HVDC bus 20. In
one proposed implementation, the FPMC 12 is a computer configured
(e.g., by software) to manage the HVDC power supplied to the
propulsion motor controllers 28.
The portion of the power supply system 10 depicted in FIG. 1
further includes two AC generators 14 and two HVDC batteries 18.
Each AC generator 14 is mechanically driven by an internal
combustion engine or a gas turbine engine (not shown in FIG. 1).
Only one AC generator 14 and one HVDC battery 18 (associated with
one wing of the hybrid electric aircraft) are shown in FIG. 1.
As used herein, the term "battery" includes least one battery cell
inside a battery case and at least one sensor located inside or
outside the battery case. In a preferred embodiment, the battery
includes a temperature sensor, a current sensor and a voltage
sensor. In the context of the active voltage controller described
in more detail below, the term "measuring the battery current"
means receiving an output of the current sensor and converting the
current sensor output to a measurement value.
Referring again to FIG. 1, the terminals (not shown in the
drawings) of the HVDC battery 18 are directly electrically
connected by electrical wiring to the HVDC bus 20. The terminals of
the AC generator 14 are electrically connected by electrical wiring
to an AC bus 16 via a contactor 2a.
As used herein, the term "contactor" means an electrically
controlled switch. Typically the contactor is controlled by a
circuit which has a lower power level than the switched circuit. A
typical contactor comprises contacts and an electromagnet contained
in a housing. The contacts are the current-carrying parts of the
contactor. The electromagnet provides the driving force to close
the contacts. A spring may be provided to return the electromagnet
core to its open position relative to the electromagnet coil. The
housing is an enclosure made of an electrically insulating
material.
The portion of the power supply system 10 depicted in FIG. 1
further includes an active rectifier unit 40 having input terminals
that are electrically connected by electrical wiring to the AC bus
16 via a contactor 2b. The output terminals of the active rectifier
unit 40 are also electrically connected by electrical wiring to the
HVDC bus 20. The active rectifier unit 40 is controlled to operate
in a manner that incoming AC power received from the AC bus 16 is
converted to outgoing HVDC power supplied to the HVDC bus 20. The
propulsion motor controllers 28 and multiple other non-propulsive
HVDC loads 30 receive HVDC power from the HVDC bus 20.
The portion of the power supply system 10 depicted in FIG. 1
further includes a regulated transformer rectifier unit (RTRU) 32
having input terminals that are electrically connected by
electrical wiring to the AC bus 16 via a contactor 2c. The system
also includes a low-voltage direct-current (LVDC) bus 34 that
receives regulated AC power from the regulated transformer
rectifier unit 32. In addition, other non-propulsive AC loads 38
receive AC power from the AC bus 16 via respective contactors 2d
and connecting electrical wiring.
The portion of the power supply system 10 depicted in FIG. 1
further includes an active voltage controller 22 that is configured
to receive (wirelessly or via wires) sensor data 6 representing the
state of the from the HVDC battery 18 and then send (wirelessly or
via wires) a control signal 8 representing a magnitude of an HVDC
bus reference voltage to the active rectifier unit 40. The active
rectifier unit 40 in turn controls the voltage of the HVDC bus 20
to match the HVDC bus reference voltage. If the HVDC bus voltage is
greater than the voltage of the HVDC battery 18 (hereinafter
"battery voltage"), the HVDC battery 18 charges. The charge rate is
proportional to the magnitude of the difference between the HVDC
bus reference voltage and the battery voltage. Conversely, if the
HVDC bus voltage is less than the voltage of the HVDC battery 18,
the HVDC battery 18 discharges. The discharge rate is proportional
to the magnitude of the difference between the HVDC bus reference
voltage and the battery voltage. Thus the active voltage controller
22 effectively controls the charge or discharge rate of the HVDC
battery 18 by adjusting the HVDC bus reference voltage supplied to
the active rectifier unit 40.
In accordance with one embodiment, the active voltage controller 22
is configured to receive different mission requests (some of which
are described below) from the FPMC 12. As used herein, the term
"mission request" means information carried by a control signal
sent (wirelessly or via wiring) from the FPMC 12 to the active
voltage controller 22 (as indicated by a dashed arrow 4 in FIG. 1),
which information identifies a change in the HVDC bus voltage to be
executed by the active rectifier unit 40.
In accordance with some embodiments, the electric generator 14 is a
wound rotor generator controlled by a generator controller unit
(not shown in FIG. 1). The wound rotor generator voltage magnitude
is controlled by the generator controller unit varying excitation
in the rotor field winding. The frequency of the AC bus voltage
cannot be controlled. Because of that, AC loads have a motor
controller or a power converter at their front end to adjust to the
varying nature of the AC bus power. The power converter that
converts AC bus power to HVDC bus power is the active rectifier
unit 40. The active rectifier unit 40 independently controls the
HVDC bus voltage. Because the HVDC battery 18 is directly connected
to the HVDC bus 20, the active voltage controller 22 is provided to
control the battery charge or discharge rate and for other
high-level control techniques to optimize energy usage on the
hybrid electric aircraft.
In accordance with other embodiments, the generator 14 is a
permanent magnet generator (not shown in FIG. 1). A permanent
magnet generator is a generator wherein the excitation field is
provided by a rotating permanent magnet instead of a coil. The
permanent magnet generator voltage, which cannot be controlled,
changes depending on the speed of the prime mover and generator
loading. Therefore, the AC bus voltage will be variable frequency,
variable magnitude voltage. Because of that, AC loads will also
require a motor controller or a power converter at the front end,
similar to the previous case where the generator is a wound rotor
generator with a generator controller unit. Similarly, active
voltage control is provided for the power converter that converts
AC bus power to HVDC bus power because the battery charge and
discharge rates need to be controlled. In addition, the HVDC bus
active voltage controller provides extra benefits of enabling the
energy optimization control. [As used herein, the term "HVDC bus
active voltage controller" refers to the active voltage controller
22 and the active rectifier unit 40 in combination.]
The active rectifier unit 40 is a solid state power converter that
is configured to control the HVDC bus voltage using a PWM control
technique. There are many topologies for active rectifiers
including boost, buck, Vienna and others. The active rectifier unit
40 in turn is controlled by the active voltage controller 22 to
enable controlling the charge and discharge rates of the HVDC
battery 18 and enable energy use optimization on the hybrid
electric aircraft.
FIG. 2 is a block diagram identifying some components of the
solid-state active rectifier unit 40 included in the architecture
depicted in FIG. 1. In accordance with one embodiment, the active
rectifier unit 40 includes an active rectifier 46 controlled by an
active rectifier controller 52. The active rectifier unit 40
further includes electromagnetic interference (EMI) and harmonics
filters 44 (hereinafter "EMI and harmonics filters 44") disposed
between the active rectifier 46 and the electric generator 14 and a
filter 48 disposed between the active rectifier 46 and the HVDC bus
20 (not shown in FIG. 2). In addition, the active rectifier unit 40
includes voltage and current sensing circuits 42 disposed between
the electric generator 14 and the EMI and harmonics filters 44 as
well as voltage and current sensing circuits 50 disposed between
the filter 48 and the HVDC bus 20.
The active rectifier controller 52 is configured to execute control
algorithms that control the states of the switches incorporated in
the active rectifier 46. The typical control algorithms for the
active rectifier unit 40 include power factor correction (PFC)
control for meeting AC bus power quality requirements and HVDC bus
control for HVDC bus voltage regulation and HVDC bus current
limitation and fault protection. The voltage and current sensing
circuits 42 sense the voltage V.sub.ac and current I.sub.ac at the
AC side of the active rectifier 46. The outputs of the voltage and
current sensing circuits 42 are received by the active rectifier
controller 52 and used to enable PFC control. The voltage and
current sensing circuits 50 sense the voltage V.sub.dc and current
I.sub.dc at the DC side of the active rectifier 46 are received by
the active rectifier controller 52 and used to enable HVDC bus
voltage regulation and current limitation and fault protection. The
EMI and harmonics filters 44 and 48 (e.g., low-pass filters, each
filter comprising an inductor and a capacitor) on the AC and HVDC
sides of the active rectifier 46 are needed for meeting power
quality requirements. The active rectifier 46 consists of a switch
network that is controlled using PWM to reduce the DC component of
voltage.
Pulse width modulation is a modulation technique that can be used
to control the power supplied to electrical devices. The average
value of voltage (and current) fed to the load is controlled by
turning the switch between supply and load on and off at a fast
rate. The longer the switch is on compared to the off periods, the
higher the total power supplied to the load. The term "duty cycle"
describes the proportion of ON time to the regular interval or
"period" of time; a low duty cycle corresponds to low power,
because the power is off for most of the time. The main advantage
of PWM is that power loss in the switching devices is very low.
When a switch is off there is practically no current, and when it
is on and power is being transferred to the load, there is almost
no voltage drop across the switch. Power loss, being the product of
voltage and current, is thus in both cases close to zero.
In accordance with the embodiments disclosed herein, in addition to
the typical active rectifier controls shown in FIG. 2, an extra
outer feedback loop is provided to enable control of the HVDC
battery charge and discharge rates. Additional control loops
managing power flow between the battery and the generator source
are also possible. Power flow management can be used to optimize
energy use efficiency depending on various conditions such as
battery SOC, remaining hydrocarbon fuel for the generator prime
mover, flight phase, aircraft mission, etc. One example control
diagram for the active voltage controller 22 that allows for
battery charge and discharge rate control and for power flow
management is shown on FIG. 3. The output of the active voltage
controller 22 to the active rectifier unit 40 is an HVDC bus
reference voltage which is either greater or less than the battery
voltage. The active voltage controller 22 calculates a new HVDC bus
reference voltage (whether it needs to be decreases or increased)
and communicates new HVDC bus reference voltage to the active
rectifier unit 40. More specifically, the active voltage controller
22 includes a current sensor that measures the battery current
(hereinafter "measured battery current"). The active voltage
controller 22 then controls the HVDC bus voltage by setting and
sending a HVDC bus reference voltage to the active rectifier unit
40, which HVDC bus reference voltage is a function of the measured
battery current.
FIG. 3 is a flowchart identifying steps of a control algorithm 100
performed by the active voltage controller 22 in accordance with
one proposed implementation. The active voltage controller 22
determines the HVDC bus reference voltage in blocks "Increase HVDC
bus voltage" (step 106) and "Decrease HVDC bus voltage" (step 110)
shown in FIG. 3. The outcome of execution of steps 106 and 110 by
the active rectifier unit 40 will be a new HVDC bus reference
voltage.
First, the active voltage controller 22 measures the HVDC battery
current (step 102). Then the active voltage controller 22
determines whether the battery current is greater than the maximum
discharge rate or not (step 104). If a determination is made in
step 104 that the battery current is discharging at a discharge
rate greater than the maximum discharge rate, then the active
voltage controller 22 increases the HVDC bus reference voltage and
communicates that value to the active rectifier unit 40 by sending
a control signal representing a magnitude of the new HVDC bus
reference voltage. The active rectifier unit 40 responds to that
control signal by increasing the HVDC bus voltage (step 106),
thereby reducing the battery discharge rate. If a determination is
made in step 104 that the battery current is not greater than the
maximum discharge rate, then the active voltage controller 22
determines whether the current into the battery is greater than the
maximum charge rate or not (step 108). If a determination is made
in step 108 that the battery current is charging at a charge rate
greater than the maximum charge rate, then the active voltage
controller 22 decreases the HVDC bus reference voltage and
communicates that value to the active rectifier unit 40 by sending
a control signal representing a magnitude of the new HVDC bus
reference voltage. The active rectifier unit 40 responds to that
control signal by decreasing the HVDC bus voltage (step 110),
thereby reducing the battery charge rate. If a determination is
made in step 108 that the current into the battery is not greater
than the maximum charge rate, then the active voltage controller 22
next determines whether the battery state of charge (SOC) is less
than the minimum SOC reserve or not (step 112).
State of charge (SOC) is the equivalent of a fuel gauge for the
battery pack in a battery-powered electric vehicle. The units of
SOC are percentage points (0%=empty; 100%=full). Usually, the SOC
cannot be measured directly but can be estimated from direct
measurement variables. For example, the active voltage controller
22 may be configured to calculate the SOC by integrating the
battery current over time.
If a determination is made in step 112 that the battery SOC is not
less than the minimum SOC reserve, then the active voltage
controller 22 determines whether a mission request to increase the
power drawn from the battery has been received or not (step 114).
If a determination is made in step 114 that a mission request to
increase the power drawn from the battery has not been received,
then the active voltage controller 22 returns to step 102. If a
determination is made in step 114 that a mission request to
increase the power drawn from the battery has been received, then
the active voltage controller 22 decreases the HVDC bus reference
voltage and communicates that value to the active rectifier unit 40
by sending a control signal representing a magnitude of the new
HVDC bus reference voltage. The active rectifier unit 40 responds
to that control signal by decreasing the HVDC bus voltage (step
110), thereby increasing the discharge rate of the battery (step
116). The active voltage controller 22 then returns to step
102.
If a determination is made in step 112 that the battery SOC is less
than the minimum SOC reserve, then the active voltage controller 22
determines whether the fuel reserve is less than the minimum fuel
reserve or not (step 118). If a determination is made in step 118
that the fuel reserve is less than the minimum fuel reserve, then
the active voltage controller 22 sends a message to the pilots to
initiate descent (step 120). If a determination is made in step 118
that the fuel reserve is not less than the minimum fuel reserve,
then the active voltage controller 22 performs steps 122 and
126.
In step 122, the active voltage controller 22 determines whether
the whether a mission request to increase the power drawn from the
generator has been received or not. If a determination is made in
step 122 that a mission request to increase the power drawn from
the generator has not been received, then the active voltage
controller 22 returns to step 102. If a determination is made in
step 122 that a mission request to increase the power drawn from
the generator has been received, then the active voltage controller
22 increases the HVDC bus reference voltage and communicates that
value to the active rectifier unit 40 by sending a control signal
representing a magnitude of the new HVDC bus reference voltage. The
active rectifier unit 40 responds to that control signal by
increasing the HVDC bus voltage (step 124), thereby reducing the
battery discharge rate. (As used herein, the terms "decrease" and
"reduce" are interchangeable synonyms.) The active voltage
controller 22 then returns to step 102.
In step 126, the active voltage controller 22 determines whether
the whether a mission request to charge the battery has been
received or not. If a determination is made in step 126 that a
mission request to charge the battery has not been received, then
the active voltage controller 22 returns to step 102. If a
determination is made in step 126 that a mission request to charge
the battery has been received, then the active voltage controller
22 increases the HVDC bus reference voltage and communicates that
value to the active rectifier unit 40 by sending a control signal
representing a magnitude of the new HVDC bus reference voltage. The
active rectifier unit 40 responds to that control signal by
increasing the HVDC bus voltage to a level greater than the battery
voltage (step 128), thereby first reducing the discharge rate and
then increasing the charge rate of the battery. The active voltage
controller 22 then returns to step 102.
The active voltage controller 22 also receives commands from the
aircraft flight power management controller 12 (hereinafter "FPMC
12"). The FPMC 12 is a computer configured (e.g., by software) to
direct the active voltage controller 22 to draw more power from
HVDC battery 18 or from the electric generator 14 or to charge the
HVDC battery 18. The FPMC 12 also receives information from the
active voltage controller 22 on status of the HVDC battery 18, such
as current battery SOC and current charging or discharging rate.
Based on mission profile, aircraft status, phase of flight, etc.,
the FPMC 12 is configured to make decisions on how to manage power
flow in the system.
Aircraft missions vary from one flight to another. Some flights are
short range and some have longer ranges. Aircraft weight also could
change from one flight to another, which weight change affects the
range of the aircraft. For very short flights and/or lighter loads,
it would be advantageous to use more battery power to reduce fuel
consumption. Typically the battery or batteries onboard the
aircraft are recharged on the ground. Conversely, for longer
flights and heavier loads, more hydrocarbon fuel could be used,
while the battery provides assistance during some flight phases.
During descent, if electric propulsion motors are used to
decelerate an aircraft, the resulting regenerative power could be
used to charge the battery. The function of the FPMC 12 is to
perform power flow management in a hybrid electric aircraft to
optimize overall energy usage and reduce fuel burn during a
particular mission.
FIG. 4 is a flowchart identifying steps of a control algorithm 140
performed by the FPMC 12 in accordance with one proposed
implementation. The FPMC 12 is configured to make a determination
whether the aircraft is on the ground or not (step 142). On the one
hand, if a determination is made in step 142 that the aircraft is
on the ground, then the FPMC 12 determines whether the battery is
discharged or not (step 144). On the other hand, if a determination
is made in step 142 that the aircraft is not on the ground, then
the FPMC 12 determines what the current phase of the flight of the
aircraft is (step 162).
Regardless of the outcome of step 142, the FPMC 12 determines in
step 154 whether a mission profile has been input by the crew or
not. On the one hand, if a determination is made in step 154 that a
mission profile has not been input by the crew, the FPMC 12 loads a
default battery usage profile (step 156). The default battery usage
profile can be established by the typical airplane mission (short-
or long-range mission), weight of the aircraft or number of people
and amount of cargo on board, and optimization with respect to fuel
burn reduction. On the other hand, if a determination is made in
step 154 that a mission profile has been input by the crew, the
FPMC 12 then determines the battery SOC and the fuel level (step
158). Then the FPMC 12 loads an optimum fuel usage battery profile
for a mission (step 160).
The optimum fuel usage battery profile is configured so that the
battery is utilized to minimize conventional fuel burn and maximize
battery usage. Depending on battery SOC and remaining electrical
energy in the battery, weight of the airplane and airplane mission,
the optimum fuel usage battery profile will minimize fuel burn by
maximizing use of the battery power throughout the flight cycle.
The optimum fuel usage battery profile will also ensure that the
battery is not over-discharged at the end of flight cycle and that
the battery discharge rate does not exceed maximum allowable
discharge rate during flight. The default battery usage profile may
be different than the optimum fuel usage battery profile. For
example, the default battery usage profile may not be optimal from
the fuel burn standpoint, but may be optimal from the standpoint of
speed or other mission parameter.
As previously mentioned, step 144 is performed following a
determination by the FPMC 12 that the aircraft is on the ground. On
the one hand, if a determination is made in step 144 that the
battery is not discharged, then the FPMC 12 returns to step 142. On
the other hand, if a determination is made in step 144 that the
battery is discharged, then the FPMC 12 determines whether any one
of a plurality of power sources (e.g., external power, auxiliary
power unit or engine power) is available or not (step 146). On the
one hand, if a determination is made in step 146 that a power
source is not available, then the FPMC 12 returns to step 142. On
the other hand, if a determination is made in step 146 that a power
source is available, then the FPMC 12 determines how much power can
be used for charging the HVDC battery 18. After determining how
much power can be used for battery charging, the FPMC 12 sets the
battery charging current (step 150) and then commands the active
voltage controller 22 to charge the battery using the set battery
charging current (step 152).
As previously mentioned, step 162 is performed following a
determination by the FPMC 12 that the aircraft is not on the
ground. Following determination of the aircraft flight phase, the
FPMC 12 next determines the battery SOC and the fuel level (step
164). Then the FPMC 12 adjusts the battery usage profile in
dependence on the flight phase, battery SOC and fuel level (step
166). After adjusting the battery usage profile, the FPMC 12 sets
the battery discharging or charging current in accordance with the
adjusted battery usage profile (step 168) and then commands the
active voltage controller 22 to discharge or charge the battery
using the set battery discharging or charging current.
While systems and methods for actively controlling the voltage of a
bus that supplies electric power from a battery and/or generator to
propulsion motors of a hybrid electric aircraft have been described
with reference to various embodiments, it will be understood by
those skilled in the art that various changes may be made and
equivalents may be substituted for elements thereof without
departing from the scope of the teachings herein. In addition, many
modifications may be made to adapt the teachings herein to a
particular situation without departing from the scope thereof.
Therefore it is intended that the claims not be limited to the
particular embodiments disclosed herein.
The embodiments disclosed above use one or more computer systems.
As used in the claims, the term "computer system" comprises a
single processing or computing device or multiple processing or
computing devices that communicate via wireline or wireless
connections. Such processing or computing devices typically include
one or more of the following: a processor, a controller, a central
processing unit, a microcontroller, a reduced instruction set
computer processor, an application-specific integrated circuit, a
programmable logic circuit, a field-programmable gated array, a
digital signal processor, and/or any other circuit or processing
device capable of executing the functions described herein. The
above examples are exemplary only, and thus are not intended to
limit in any way the definition and/or meaning of the term
"computer system".
The methods described herein may be encoded as executable
instructions embodied in a non-transitory tangible
computer-readable storage medium, including, without limitation, a
storage device and/or a memory device. Such instructions, when
executed by a processing or computing system, cause the system
device to perform at least a portion of the methods described
herein.
The process claims set forth hereinafter should not be construed to
require that the steps recited therein be performed in alphabetical
order (any alphabetical ordering in the claims is used solely for
the purpose of referencing previously recited steps) or in the
order in which they are recited unless the claim language
explicitly specifies or states conditions indicating a particular
order in which some or all of those steps are performed. Nor should
the process claims be construed to exclude any portions of two or
more steps being performed concurrently or alternatingly unless the
claim language explicitly states a condition that precludes such an
interpretation.
The structure corresponding to the term "switching means" recited
in the appended claims includes contacts, relays and structural
equivalents thereof.
* * * * *